1. Field of the Invention
The field of the invention is related to methods and systems for predicting material fatigue and damage. In particular the invention relates to an improved method and system for applying an improved model of material fatigue and damage for a component design, maintenance and its life prediction.
2. Description of the Related Technology
The construction and maintenance of aircraft is a complex undertaking that can result in catastrophic effects if performed improperly. During the manufacture and usage of structural components of aircraft various defects can occur. Defects can occur during the manufacture and fabrication of components. Defects can also occur during the assembly, repair and maintenance of the structural components. These defects may be latent in the structural material used in the construction of a component and may therefore exist well before actual fabrication of the component. Pre-existing defects can result in cracks being initiated under a variety of service loading and environmental conditions. Such defects can occur well below the non-destructive (NDE) inspection limits, and grow with service, finally leading to failure of the components.
Depending on the function and usage of aircraft components, the failure impact on the overall system can vary from a minor degradation to catastrophic failure. Histories of aircraft failures have indicated that failure causes are mostly initiated due to: (1) stress concentration sites such as such as rivet holes, lap joints, wing root areas etc.; (2) defects introduced during inspection, repair and assembly; or (3) initial manufacturing defects. These causal relationships have had a major effect on: (1) design philosophy, (2) the selection and use of materials with low incidences of defects, and (3) the introduction of FAA enforced periodic inspection and maintenance procedures.
Over the past four decades, there have been about 70 different crack initiation and about 40 crack growth empirical models proposed. Since the early 1970's, these models have been presented in terms of computer algorithms for predicting the life of components. There are several computer models that have been proposed by engineers in various countries to predict the life of a component using service load data. These models have not succeeded in making consistently reliable predictions. A few examples of these models are: CORPUS (from Greece), PREFAS (from Portugal), ONERA (from France), MODGRO (from US Air Force) and FASTRAN (from US NASA). FASTRAN is sometime labeled as NASTRAN.
In FIG. 1, the life predictions from the above mentioned models for an aircraft spectrum under a flight-by-flight load history for 2024 alloy (using as a mean load of 75 MPa) are compared. All of these models tend to under predict the flights to failure, with the FASTRAN model most closely approximating the test data. The models can also over predict the flights to failure. Additional details regarding FIG. 1 can be found in Lazzeri L, Pieracci A, Salvetti A; 18th Symposium of International Committee on Aeronautical Fatigue, Melbourne (Australia), May 1995.
Commonly, the inadequacies of the predictive abilities of the models are compensated by using several adjustable parameters which are correlated and tuned using lab test data. Due to the uncertainty in predictive capabilities, vehicle safety is guarded using a variety of methods. These methods may involve the use of safety factors in design, the selected use of component data, conducting periodic non-destructive (NDE) inspections, the use of statistics to assign data scatter and institution of material quality control procedures.
Ultimately the drawbacks in the current fatigue life prediction methods stem from several sources: (1) the assumption of plasticity induced crack closure; (2) the lack of terms in the models that relate to the environmental effects and slip deformation behavior of materials; (3) the requirement of several adjustable parameters that are needed in order to fit the observed data; and (4) the need for extensive lab test data to determine the adjustable parameters. Furthermore, past models use only one driving force parameter, ΔK, which is based on the “crack closure” model. Such models are not applicable to different types of service load-history and they are not applicable to all types of materials and platforms. They also cannot reliably predict component response to compression dominated service loads or predict crack initiation. Due to these drawbacks, the lack of reliability and predictability of past models stem from inaccurate and inadequate accounting of fatigue damage forcing them to be mere curve-fitting models.
Additional aspects of the invention will be apparent from the following summary, drawing figures, and detailed description.